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CHAPTER-1
BASIC MECHANISM
1.1 INTRODUCTION
The development of the gas turbine engine as an aircraft power plant has been so rapid
that it is difficult to appreciate that prior to the 1950s very few people had heard of this
method of aircraft propulsion. The possibility of using a reaction jet had interested
aircraft designers for a long time, but initially the low speeds of early aircraft and the
unsuitably of a piston engine for producing the large high velocity airflow necessary for
the ‘jet’ presented many obstacles.
A French engineer, René Lorin, patented a jet propulsion engine (fig. 1.1) in 1913, but
this was an athodyd and was at that period impossible to manufacture or use, since
suitable heat resisting materials had not then been developed and, in the second place, jet
propulsion would have been extremely inefficient at the low speeds of the aircraft of
those days. However, today the modern ram jet is very similar to Lorin's conception.
In 1930 Frank Whittle was granted his first patent for using a gas turbine to produce a
propulsive jet, but it was eleven years before his engine completed its first flight. The
Whittle engine formed the basis of the modern gas turbine engine, and from it was
developed the Rolls-Royce Welland, Derwent, Nene and Dart engines.
The Derwent and Nene turbo-jet engines had world-wide military applications; the Dart
turbo-propeller engine became world famous as the power plant for the Vickers Viscount
aircraft.
Although other aircraft may be fitted; with later engines termed twin-spool, triple-spool,
by-pass, ducted fan, unducted fan and propfan, these are inevitable developments of
Whittle's early engine.
Fig. 1.1 Lorin's jet engine. Fig. 1.2 Whittle-type turbo-jet engine.
The jet engine (fig. 1.2), although appearing so different from the piston engine-propeller
combination, applies the same basic principles to effect propulsion. As shown in fig. 1.3,
both propel their aircraft solely by thrusting a large weight of air backwards. Although
today jet propulsion is popularly linked with the gas turbine engine, there are other types
of jet propelled engines, such as the ram jet, the pulse jet, the rocket, the turbo/ram jet,
and the turbo-rocket.
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1.2 PRINCIPLES OF JET PROPULSION
Jet propulsion is a practical application of Sir Isaac Newton's third law of motion which
states that, 'for every force acting on a body there is an opposite and equal reaction'. For
aircraft propulsion, the 'body' is atmospheric air that is caused to accelerate as it passes
through the engine. The force required to give this acceleration has an equal effect in the
opposite direction acting on the apparatus producing the acceleration. A jet engine
produces thrust in a similar way to the engine/propeller combination. Both propel the
aircraft by thrusting a large weight of air backwards (fig. 1.3), one in the form of a large
air slipstream at comparatively low speed and the other in the form of a jet of gas at very
high speed.
Fig. 1.3 Jet propulsion and propeller.
This same principle of reaction occurs in all forms of movement and has been usefully
applied in many ways. The earliest known example of jet reaction is that of Hero's engine
(fig. 1.4) produced as a toy in B.C. This toy showed how the momentum of steam issuing
from a number of jets could impart an equal and opposite reaction to the jets themselves,
thus causing the engine to revolve.
The familiar whirling garden sprinkler (fig. 1.5) is a more practical example of this
principle, for the mechanism rotates by virtue of the reaction to the water jets. The high
pressure jets of modern fire-fighting equipment are an example of 'jet reaction', for often,
due to the reaction of the water jet, the hose cannot be held or controlled by one fireman.
Perhaps the simplest illustration of this principle is afforded by the carnival balloon
which, when the air or gas is released, rushes rapidly away in the direction opposite to the
jet.
Jet reaction is definitely an internal phenomenon and does not, as is frequently assumed,
result from the pressure of the jet on the atmosphere. In fact, the jet propulsion engine,
whether rocket, athodyd, or turbo-jet, is a piece of apparatus designed to accelerate a
stream of air or gas and to expel it at high velocity. There are, of course, a number of
ways of doing this, but in all instances the resultant reaction or thrust exerted on the
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engine is proportional to the mass or weight of air expelled by the engine and to the
velocity change imparted to it. In other words, the same thrust can be provided either by
giving a large mass of air a little extra velocity or a small mass of air a large extra
velocity. In practice the former is preferred, since by lowering the jet velocity relative to
the atmosphere a higher propulsive efficiency is obtained.
Fig. 1.4 Heroís engine – probably Fig. 1.5 A garden sprinkler rotated
the earliest form of jet reaction. by the reaction of the water jets.
1.3 METHODS OF JET PROPULSION
The types of jet engine, whether ram jet, pulse jet, rocket, gas turbine, turbo/ram jet or
turbo-rocket, differ only in the way in which the 'thrust provider', or engine, supplies and
converts the energy into power for flight.
Fig. 1.6 A ram Jet engine. Fig. 1.7 A pulse jet engine.
The ram jet engine (fig. 1.6) is an athodyd, or 'aero-thermodynamic-duct to give it its full
name. It has no major rotating parts and consists of a duct with a divergent entry and a
convergent or convergent-divergent exit. When forward motion is imparted to it from an
external source, air is forced into the air intake where it loses velocity or kinetic energy
and increases its pressure energy as it passes through the diverging duct. The total energy
is then increased by the combustion of fuel, and the expanding gases accelerate to
atmosphere through the outlet duct. A ram jet is often the power plant for missiles and
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target vehicles; but is unsuitable as an aircraft power plant "because it requires forward
motion imparting to it before any thrust is produced.
Fig. 1.8 Mechanical arrangement of gas turbine engines.
The pulse jet engine (fig. 1.7) uses the principle of intermittent combustion and unlike the
ram jet it can be run at a static condition. The engine is formed by an aerodynamic duct
similar to the ram jet but, due to the higher pressures involved; it is of more robust
construction. The duct inlet has a series of inlet 'valves' that are spring-loaded into the
open position. Air drawn through the open valves passes into the combustion chamber
and is heated by the burning of fuel injected into the chamber. The resulting expansion
causes a rise in pressure, forcing the valves to close, and the expanding gases are then
ejected rearwards. A depression created by the exhausting gases allows the valves to open
and repeat the cycle. Pulse jets have been designed for helicopter rotor propulsion and
some dispense with inlet valves by careful design of the ducting to control the changing
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pressures of the resonating cycle. The pulse jet is unsuitable as an aircraft power plant
because it has high fuel consumption and is unable to equal the performance of the
modern gas turbine engine. Although a rocket engine (fig. 1.8) is a jet engine, it has one
major difference in that it does not use atmospheric air as the propulsive fluid stream.
Instead, it produces its own propelling fluid by the combustion of liquid or chemically
decomposed fuel with oxygen, which it carries, thus enabling it to operate outside the
earth's atmosphere. It is, therefore, only suitable for operation over short periods.
Fig. 1.9 Mechanical arrangement of gas turbine engines.
The application of the gas turbine to jet propulsion has avoided the inherent weakness of
the rocket and the athodyd, for by the introduction of a turbine-driven compressor a
means of producing thrust at low speeds is provided. The turbo-jet engine draws air from
the atmosphere and after compressing and heating it, a process that occurs in all heat
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engines, the energy and momentum given to the air forces it out of the propelling nozzle
at a velocity of up to 2,000 feet per second or about 1,400 miles per hour. On its way
through the engine, the air gives up some of its energy and momentum to drive the
turbine that powers the compressor.
Fig. 1.10 Comparative propulsive efficiencies.
The mechanical arrangement of the gas turbine engine is simple, for it consists of only
two main rotating parts, a compressor and a turbine, and one or a number of combustion
chambers. The mechanical arrangement of various gas turbine engines is shown in fig. 1
.9. This simplicity, however, does not apply to all aspects of the engine, for as described
in subsequent Parts the thermo and aerodynamic problems are somewhat complex. They
result from the high operating temperatures of the combustion chamber and turbine, the
effects of varying flows across the compressor and turbine blades, and the design of the
exhaust system through which the gases are ejected to form the propulsive jet.
At aircraft speeds below approximately 450 miles per hour, the pure jet engine is less
efficient than a propeller-type engine, since its propulsive efficiency depends largely on
its forward speed; the pure turbo-jet engine is, therefore, most suitable for high forward
speeds. The propeller efficiency does, however, decrease rapidly above 350 miles per
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hour due to the disturbance of the airflow caused by the high blade-tip speeds of the
propeller. These characteristics have led to some departure from the use of pure turbo-jet
propulsion where aircraft operate at medium speeds by the introduction of a combination
of propeller and gas turbine engine.
The advantages of the propeller/turbine combination have to some extent been offset by
the introduction of the by-pass, ducted fan and propfan engines. These engines deal with
larger comparative airflows and lower jet velocities than the pure jet engine, thus giving a
propulsive efficiency which is comparable to that of the turbo-prop and exceeds that of
the pure jet engine (fig. 1.10).
The turbo/ram jet engine (fig. 1.11) combines the turbo-jet engine (which is used for
speeds up to Mach 3) with the ram jet engine, which has good performance at high Mach
numbers.
Fig. 1.11 A turbo/ram jet engine.
The engine is surrounded by a duct that has a variable intake at the front and an
afterburning jet pipe with a variable nozzle at the rear. During take-off and acceleration,
the engine functions as a conventional turbo-jet with the afterburner lit; at other flight
conditions up to Mach 3, the afterburner is inoperative. As the aircraft accelerates through
Mach 3, the turbo-jet is shut down and the intake air is diverted from the compressor, by
guide vanes, and ducted straight into the afterburning jet pipe, which becomes a ram jet
combustion chamber. This engine is suitable for an aircraft requiring high speed and
sustained high Mach number cruise conditions where the engine operates in the ram jet
mode. The turbo-rocket engine (fig. 1.12) could be considered as an alternative engine to
the turbo/ram jet; however, it has one major difference in that it carries its own oxygen to
provide combustion.
The engine has a low pressure compressor driven by a multi-stage turbine; the power to
drive the turbine is derived from combustion of kerosene and liquid oxygen in a rocket-
type combustion chamber. Since the gas temperature will be in the order of 3,500 deg. C,
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additional fuel is sprayed into the combustion chamber for cooling purposes before the
gas enters the turbine. This fuel-rich mixture (gas) is then diluted with air from the
compressor and the surplus fuel burnt in a conventional afterburning system.
Fig. 1.12 A turbo-rocket engine.
Although the engine is smaller and lighter than the turbo/ram jet, it has a higher fuel
consumption. This tends to make it more suitable for an interceptor or space-launcher
type of aircraft that requires high speed, high altitude performance and normally has a
flight plan that is entirely accelerative and of short duration.
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CHAPTER- 2
WORKING CYCLE AND AIR FLOW
2.1 INTRODUCTION The gas turbine engine is essentially a heat engine using air as a working fluid to provide
thrust. To achieve this, the air passing through the engine has to be accelerated; this
means that the velocity or kinetic energy of the air is increased. To obtain this increase,
the pressure energy is first of all increased, followed by the addition of heat energy,
before final conversion back to kinetic Energy in the form of a high velocity jet efflux.
2.2 WORKING CYCLE
The working cycle of the gas turbine engine is similar to that of the four-stroke piston
engine. However, in the gas turbine engine, combustion occurs at a constant pressure,
whereas in the piston engine it occurs at a constant volume. Both engine cycles (fig. 2-1)
show that in each instance there is induction, compression, combustion and exhaust.
These processes are intermittent in the case of the piston engine whilst they occur
continuously in the gas turbine. In the piston engine only one stroke is utilized in the
production of power, the others being involved in the charging, compressing and
exhausting of the working fluid. In contrast, the turbine engine eliminates the three 'idle'
strokes, thus enabling more fuel to be burnt in a shorter time; hence it produces a greater
power output for a given size of engine.
Due to the continuous action of the turbine engine and the fact that the combustion
chamber is not an enclosed space, the pressure of the air does not rise, like that of the
piston engine, during combustion but its volume does increase. This process is known as
heating at constant pressure. Under these conditions there is no peak or fluctuating
pressures to be withstood, as is the case with the piston engine with its peak pressures in
excess of 1,000 lb. per sq. in. It is these peak pressures which make it necessary for the
piston engine to employ cylinders of heavy construction and to use high octane fuels, in
contrast to the low octane fuels and the light fabricated combustion chambers used on the
turbine engine.
The working cycle upon which the gas turbine engine functions is, in its simplest form,
represented by the cycle shown on the pressure volume diagram in fig. 2.2. Point A
represents air at atmospheric pressure that is compressed along the line AB. From B to C
heat is added to the air by introducing and burning fuel at constant pressure, thereby
considerably increasing the volume of air. Pressure losses in the combustion chambers
(Part 4) are indicated by the drop between B and C. From C to D the gases resulting from
combustion expand through the turbine and jet pipe back to atmosphere. During this part
of the cycle, some of the energy in the expanding gases is turned into mechanical power
by the turbine; the remainder, on its discharge to atmosphere, provides a propulsive jet.
Because the turbo-jet engine is a heat engine, the higher the temperature of combustion
the greater is the expansion of the gases. Under these conditions there is no peak or
fluctuating pressures to be withstood, as is the case with the piston engine with its peak
pressures in excess of 1,000 lb. per sq. in. The combustion temperature, however, must
not exceed a value that gives a turbine gas entry temperature suitable for the design and
materials of the turbine assembly.
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Fig. 2.1 A comparison between the working cycle of a turbo-jet engine and a piston
engine.
The use of air-cooled blades in the turbine assembly permits a higher gas temperature and
a consequently higher thermal efficiency.
Fig. 2.2 The working cycle on a pressure-volume diagram.
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2.3 THE RELATIONS BETWEEN PRESSURE, VOLUME AND
TEMPERATURE
During the working cycle of the turbine engine, the airflow or 'working fluid' receives
and gives up heat, so producing changes in its pressure, volume and temperature. These
changes as they occur are closely related, for they follow a
Fig. 2.3 An airflow through divergent and convergent ducts.
common principle that is embodied in a combination of the laws of Boyle and Charles.
Briefly, this means that the product of the pressure and the volume of the air at the
various stages in the working cycle is proportion-al to the absolute temperature of the air
at those stages. This relationship applies for whatever means are used to change the state
of the air. For example, whether energy is added by combustion or by compression, or is
extracted by the turbine, the heat change is directly proportional to the work added or
taken from the gas.
There are three main conditions in the engine working cycle during which these changes
occur. During compression, when work is done to increase the pressure and decrease the
volume of the air, there is a corresponding rise in the temperature. During combustion,
when fuel is added to the air and burnt to increase the temperature, there is a
corresponding increase in volume whilst the pressure remains almost constant. During
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expansion, when work is taken from the gas stream by the turbine assembly, there is a
decrease in temperature and pressure with a corresponding increase in volume. Changes
in the temperature and pressure of the air can be traced through an engine by using the
airflow diagram in fig. 2.5. With the airflow being continuous, volume changes are
shown up as changes in velocity.
The efficiency with which these changes are made will determine to what extent the
desired relations between the pressure, volume and temperature are attained. For the more
efficient the compressor, the higher the pressure generated for a given work input; that is,
for a given temperature rise of the air. Conversely, the more efficiently the turbine uses
the expanding gas, the greater the output of work for a given pressure drop in the gas.
Changes in the temperature and pressure of the air can be traced through an engine by
using the airflow diagram in fig. 2.5. With the airflow being continuous, volume changes
are shown up as changes in velocity.
When the air is compressed or expanded at 100 per cent efficiency, the process is said to
be adiabatic. Since such a change means there is no energy losses in the process, either by
friction, conduction or turbulence, it is obviously impossible to achieve in practice; 90 per
cent is a good adiabatic efficiency for the compressor and turbine.
2.4 CHANGES IN VELOCITY AND PRESSURE
During the passage of the air through the engine, aerodynamic and energy requirements
demand changes in its velocity and pressure. For instance: during compression, a rise in
the pressure of the air is required and not an increase in its velocity. After the air has been
heated and its internal energy increased by combustion, an increase in the velocity of the
gases is necessary to force the turbine to rotate. At the propelling nozzle a high exit
velocity is required, for it is the change in the momentum of the air that provides the
thrust on the aircraft. Local decelerations of airflow are also required, as for instance, in
the combustion chambers to provide a low velocity zone for the flame to burn.
Fig. 2.4 Supersonic airflow through a convergent-divergent nozzle or venturi.
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These various changes are effected by means of the size and shape of the ducts through
which the air passes on its way through the engine. Where a conversion from velocity
(kinetic) energy to pressure is required, the passages are divergent in shape. Conversely,
where it is required to convert the energy stored in the combustion gases to velocity
energy, a convergent passage or nozzle (fig. 2.3) is used. These shapes apply to the gas
turbine engine where the airflow velocity is subsonic or sonic, i.e. at the local speed of
sound. Where supersonic speeds are encountered, such as in the propelling nozzle of the
rocket, athodyd and some jet engines (Part 6), a convergent-divergent nozzle or venturi
(fig. 2.4) is used to obtain the maximum conversion of the energy in the combustion
gases to kinetic energy.
The design of the passages and nozzles is of great importance, for upon their good design
will depend the efficiency with which the energy changes are affected. Any interference
with the smooth airflow creates a loss in efficiency and could result in component failure
due to vibration caused by eddies or turbulence of the airflow.
2.5 AIRFLOW
Fig. 2.5.1 Airflow systems.
The path of the air through a gas turbine engine varies according to the design of the
engine. A straight-through flow system (fig. 2.5) is the basic design, as it provides for an
engine with a relatively small frontal area and is also suitable for use of the by-pass
principle. In contrast, the reverse flow system gives an engine with greater frontal area,
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but with a reduced overall length. The operation, however, of all engines is similar. The
variations due to the different designs are described in the subsequent paragraphs.
The major difference of a turbo-propeller engine is the conversion of gas energy into
mechanical power to drive the propeller. Only a small amount of 'jet thrust' is available
from the exhaust system. The majority of the energy in the gas stream is absorbed by
additional turbine stages, which drive the propeller through internal shafts.
Fig, 2.5.2 Airflow systems.
As can be seen in fig. 2.5, the by-pass principle involves a division of the airflow.
Conventionally, all the air taken in is given an initial low compression and a percentage is
then ducted to by-pass, the remainder being delivered to the combustion system in the
usual manner, this principle is conducive to improved propulsive efficiency and specific
fuel consumption.
An important design feature of the by-pass engine is the by-pass ratio; that is, the ratio of
cool air by-passed through the duct to the flow of air passed through the high pressure
system. With low by-pass ratios, i.e. in the order of 1:1, the two streams are usually
mixed before being exhausted from the engine. The fan engine may be regarded as an
extension of the by-pass principle, and the requirement for high by-pass ratios of up to
5:1 is largely met by using the front fan in a twin or triple-spool configuration (on which
the fan is, in fact, the low pressure compressor) both with and without mixing of the
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airflows. Very high by-pass ratios, in the order of 15:1, are achieved using propfans.
These are a variation on the turbo-propeller theme but with advanced technology
propellers capable of operating with high efficiency at high aircraft speeds.
On some front fan engines, the by-pass airstream is ducted overboard either directly
behind the fan through short ducts or at the rear of the engine through longer ducts; hence
the term 'ducted fan'. Another, though seldom used, variation is that of the aft fan.
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CHAPTER- 3
COMPRESSOR
3.1 INTRODUCTION
In the gas turbine engine, compression of the air before expansion through the turbine is
effected by one of two basic types of compressor, one giving centrifugal flow and the
other axial flow. Both types are driven by the engine turbine and are usually coupled
direct to the turbine shaft.
3.2 THE AXIAL FLOW COMPRESSOR
Fig. 3.1 Typical axial flow compressors.
An axial flow compressor (fig. 3.1 and fig. 3.2) consists of one or more rotor assemblies
that carry blades of airfoil section. These assemblies are mounted between bearings in the
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casings which incorporate the stator vanes. The compressor is a multi-stage unit as the
amount of pressure increase by each stage is small; a stage consists of a row of rotating
blades followed by a row of stator vanes. Where several stages of compression operate in
series on one shaft it becomes necessary to vary the stator vane angle to enable the
compressor to operate effectively at speeds below the design condition. As the pressure
ratio is increased the incorporation of variable stator vanes ensures that the airflow is
directed onto the succeeding stage of rotor blades at an acceptable angle. From the front to the rear of the compressor, i.e. from the low to the high pressure end,
there is a gradual reduction of the air annulus area between the rotor shaft and the stator
casing. This is necessary to maintain a near constant air axial velocity as the density
increases through the length of the compressor. The convergence of the air annulus is
achieved by the tapering of the casing or rotor. A combination of both is also possible,
with the arrangement being influenced by manufacturing problems and other mechanical
design factors.
Fig. 3.2 Typical triple spool compressor.
A single-spool compressor (fig. 3.1) consists of one rotor assembly and stators with as
many stages as necessary to achieve the desired pressure ratio and all the airflow from
the intake passes through the compressor.
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The multi-spool compressor consists of two or more rotor assemblies, each driven by
their own turbine at an optimum speed to achieve higher pressure ratios and to give
greater operating flexibility.
Although a twin-spool compressor (fig. 3.1) can be used for a pure jet engine, it is most
suitable for the by-pass type of engine where the front or low pressure compressor is
designed to handle a larger airflow than the high pressure compressor. Only a percentage
of the air from the low pressure compressor passes into the high pressure compressor; the
remainder of the air, the by-pass flow, is ducted around the high pressure compressor.
Both flows mix in the exhaust system before passing to the propelling nozzle. This
arrangement matches the velocity of the jet nearer to the optimum requirements of the
aircraft and results in higher propulsive efficiency, hence lower fuel consumption. For
this reason, the pure jet engine where all the airflow passes through the full compression
cycle is now obsolete for all but the highest speed aircraft.
With the high by-pass ratio turbo-fan this trend is taken a stage further. The intake air
undergoes only one stage of compression in the fan before being split between the core or
gas generator system and the by-pass duct in the ratio of approximately one to five (fig.
3.2). This results in the optimum arrangement for passenger and/or transport aircraft
flying at just below the speed of sound. The fan may be coupled to the front of a number
of core compression stages (two shaft engine) or a separate shaft driven by its own
turbine (three shaft engine).
3.2.1 Principles of operation
During operation the rotor is turned at high speed by the turbine so that air is
continuously induced into the compressor, which is then accelerated by the rotating
blades and swept rearwards onto the adjacent row of stator vanes. The pressure rise
results from the energy imparted to the air in the rotor which increases the air velocity.
The air is then decelerated (diffused) in the following stator passage and the kinetic
energy translated into pressure. Stator vanes also serve to correct the deflection given to
the air by the rotor blades and to present the air at the correct angle to the next stage of
rotor blades. The last row of stator vanes usually acts as air straighteners to remove swirl
from the air prior to entry into the combustion system at a reasonably uniform axial
velocity. Changes in pressure and velocity that occur in the airflow through the
compressor are shown diagrammatically in fig. 3.3. The changes are accompanied by a
progressive increase in air temperature as the pressure increases. The pressure rise results
from the energy imparted to the air in the rotor which increases the air velocity.
Across each stage the ratio of total pressures of outgoing air and inlet air is quite small,
being between 1:1 and 1:2. The reason for the small pressure increase through each stage
is that the rate of diffusion and the deflection angle of the .blades must be limited if losses
due to air breakaway at the blades and subsequent blade stall are to be avoided. Although
the pressure ratio of each stage is small, every stage increases the exit pressure of the
stage that precedes it. So whilst this first stage of a compressor may only increase the
pressure by 3 to 4 lb. per sq. in., at the rear of a thirty to one compression system the
stage pressure rise can be up to 80 lb, per sq. in, The ability to design multi-stage axial
compressors with controlled air velocities and straight through flow, minimizes losses
and results in a high efficiency and hence low fuel consumption. This gives it a further
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advantage over the centrifugal compressor where these conditions are fundamentally not
so easily achieved.
The more the pressure ratio of a compressor is increased the more difficult it becomes to
ensure that it will operate efficiently over the full speed range. This is because the
requirement for the ratio of inlet area to exit area, at the high speed case, results in an
inlet area that becomes progressively too large relative to the exit area as the compressor
speed and hence pressure ratio is reduced. The axial velocity of the inlet air in the front
stages thus becomes low relative to the blade speed, this changes the incidence of the air
onto the blades and a condition is reached where the flow separates and the compressor
flow breaks down. Where high pressure ratios are required from a single compressor this
problem can be overcome by introducing variable stator vanes in the front stages of the
system. This corrects the incidence of air onto the rotor blades to angles which they can
tolerate. An alternative is the incorporation of interstage bleeds, where a proportion of air
after entering the compressor is removed at an intermediate stage and .dumped into the
bypass flow. While this method corrects the axial velocity through the preceding stages,
energy is wasted and incorporation of variable stators is preferred.
Fig. 3.3 Pressure and velocity changes through an axial compressor.
The fan of the high by-pass ratio turbo-fan is an example of an axial compressor which
has been optimized to meet the specific requirements of this cycle. While similar in
principle to the core compressor stage, the proportions of design are such that the inner
gas path is similar to that of the core compressor that follows it, while the tip diameter is
considerably larger. The mass flow passed by the fan is typically six times that required
by the core, the remaining five sixths by-pass the core and is expanded through its own
coaxial nozzle, or may be mixed with the flow at exit from the core in a common nozzle.
To optimize the cycle the by-pass flow has to be raised to a pressure of approximately 1.6
times the inlet pressure. This is achieved in the fan by utilizing very high tip speeds (1500
ft. per sec.) and airflow such that the by-pass section of the blades operate with a
supersonic inlet air velocity of up to Mach 1.5 at the tip. The pressure that results is
graded from a high value at the tip where relative velocities are highest to the more
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normal values of 1.3 to 1.4 at the inner radius which supercharges the core where
aerodynamic design is more akin to that of a conventional compressor stage. The
capability of this type of compressor stage achieves the cycle requirement of high flow
per unit of frontal area, high efficiency and high pressure ratio in a single rotating blade
row without inlet guide vanes within an acceptable engine diameter. Thus keeping weight
and mechanical complexity at an acceptable level.
3.2.2 Construction
The construction of the compressor centers around the rotor assembly and casings. The
rotor shaft is supported in ball and roller bearings and coupled to the turbine shaft in a
manner that allows for any slight variation of alignment. The cylindrical casing assembly
may consist of a number of cylindrical casings with a bolted axial joint between each
stage or the casing may be in two halves with a bolted centre line joint. One or other of
these construction methods is required in order that the casing can be assembled around
the rotor.
3.2.3 Rotors
In compressor designs (fig. 3.4) the rotational speed is such that a disc is required to
support the centrifugal blade load. Where a number of discs are fitted onto one shaft they
may be coupled and secured together by a mechanical fixing but generally the discs are
assembled and welded together, close to their periphery, thus forming an integral drum.
Fig. 3.4 Rotors of drum and disc construction.
Typical methods of securing rotor blades to the disc are shown in fig. 3.5, fixing may be
circumferential or axial to suit special requirements of the stage. In general the aim is to
design a securing feature that imparts the lightest possible load on the supporting disc
thus minimizing disc weight. Whilst most compressor designs have separate blades for
manufacturing and maintainability requirements, it becomes more difficult on the
smallest engines to design a practical fixing. However this may be overcome by
producing blades integral with the disc; the so called 'blisk'.
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Fig. 3.5 Methods of securing blades to disc.
3.2.4 Rotor blades
The rotor blades are of airfoil section (fig. 3.6) and usually designed to give a pressure
gradient along their length to ensure that the air maintains a reasonably uniform axial
velocity. The higher pressure towards the tip balances out the centrifugal action of the
rotor on the airstream. To obtain these conditions, it is necessary to 'twist' the blade from
root to tip to give the correct angle of incidence at each point. Air flowing through a
compressor creates two boundary layers of slow to stagnant air on the inner and outer
walls. In order to compensate for the slow air in the boundary layer a localized increase in
blade camber both at the blade tip and root has been introduced. The blade extremities
appear as if formed by bending over each corner, hence the term 'end-bend'.
3.2.5 Stator vanes
The stator vanes are again of airfoil section and are secured into the compressor casing or
into stator vane retaining rings, which are themselves secured to the casing (fig. 3.7). The
vanes are often assembled in segments in the front stages and may be shrouded at their
inner ends to minimize the vibrational effect of flow variations on the longer vanes. It is
also necessary to lock the stator vanes in such a manner that they will not rotate around
the casing.
22
Fig. 3.6 A typical rotor blade showing twisted contour.
Fig. 3.7 Methods of securing vanes to compressor casing.
3.3 OPERATING CONDITIONS
Each stage of a multi-stage compressor possesses certain airflow characteristics that are
dissimilar from those of its neighbor; thus to design a workable and efficient compressor,
the characteristics of each stage must be carefully matched. This is a relatively simple
process to implement for one set of conditions (design mass flow, pressure ratio and
rotational speed), but is much more difficult when reasonable matching is to be retained
23
with the compressor operating over a wide range of conditions such as an aircraft engine
encounters.
Fig. 3.8 Limits of stable airflow.
Fig. 3.9 Typical variable stator vanes.
If the operating conditions imposed upon the compressor blade departs too far from the
design intention, breakdown of airflow and/or aerodynamically induced vibration will
occur. These phenomena may take one of two forms; the blades may stall because the
angle of incidence of the air relative to the blade is too high (positive incidence stall) or
too low (negative incidence stall). The former is a front stage problem at low speeds and
the latter usually affects the rear stages at high speed, either can lead to blade vibration
which can induce rapid destruction. If the engine demands a pressure rise from the
compressor, which is higher than the blading can sustain, 'surge' occurs.
24
CHAPTER- 4
COMBUSTION CHAMBER
4.1 INTRODUCTION
The combustion chamber (fig. 4.1) has the difficult task of burning large quantities of
fuel, supplied through the fuel spray nozzles, with extensive volumes of air, supplied by
the compressor (Part 3), and releasing the heat in such a manner that the air is expanded
and accelerated to give a smooth stream of uniformly heated gas at all conditions required
by the turbine. This task must be accomplished with the minimum loss in pressure and
with the maximum heat release for the limited space available.
The amount of fuel added to the air will depend upon the temperature rise required.
However, the maximum temperature is limited to within the range of 850 to 1700 deg. C.
by the materials from which the turbine blades and nozzles are made. The air has already
been heated to between 200 and 550 deg. C. by the work done during compression,
giving a temperature rise requirement of 650 to 1150 deg. C. from the combustion
process. Since the gas temperature required at the turbine varies with engine thrust, and in
the case of the turbo-propeller engine upon the power required, the combustion chamber
must also be capable of maintaining stable and efficient combustion over a wide range of
engine operating conditions.
Fig. 4.1 An early combustion chamber.
25
Efficient combustion has become increasingly important because of the rapid rise in
commercial aircraft traffic and the consequent increase in atmospheric pollution, which is
seen by the general public as exhaust smoke.
4.2 COMBUSTION PROCESS
Air from the engine compressor enters the combustion chamber at a velocity up to 500
feet per second, but because at this velocity the air speed is far too high for combustion,
the first thing that the chamber must do is to diffuse it, i.e. decelerate it and raise its static
pressure. Since the speed of burning kerosene at normal mixture ratios is only a few feet
per second, any fuel lit even in the diffused air stream, which now has a velocity of about
80 feet per second, would be blown away. A region of low axial velocity has therefore to
be created in the chamber, so that the flame will remain alight throughout the range of
engine operating conditions.
In normal operation, the overall air/fuel ratio of a combustion chamber can vary between
45:1 and 130:1, However, kerosene will only burn efficiently at, or close to, a ratio of
15:1, so the fuel must be burned with only part of the air entering the chamber, in what is
called a primary combustion zone. A region of low axial velocity has therefore to be
created in the chamber, so that the flame will remain alight throughout the range of
engine operating conditions.
This is achieved by means of a flame tube (combustion liner) that has various devices for
metering the airflow distribution along the chamber.
Fig. 4.2 Apportioning the airflow.
Approximately 20 per cent of the air mass flow is taken in by the snout or entry section
(fig. 4.2). Immediately downstream of the snout are swirl vanes and a perforated flare,
through which air passes into the primary combustion zone. The swirling air induces a
flow upstream of the center of the flame tube and promotes the desired recirculation. The
air not picked up by the snout flows into the annular space between the flame tube and
the air casing.
Through the wall of the flame tube body, adjacent to the combustion zone, are a selected
number of secondary holes through which a further 20 per cent of the main flow of air
passes into the primary zone. The air from the swirl vanes and that from the secondary air
holes interacts and creates a region of low velocity recirculation. This takes the form of a
toroidal vortex, similar to a smoke ring, which has the effect of stabilizing and anchoring
the flame (fig, 4.3). The recirculating gases hasten the burning of freshly injected fuel
droplets by rapidly bringing them to ignition temperature. It is arranged that the conical
26
fuel spray from the nozzle intersects the recirculation vortex at its centre. This action,
together with the general turbulence in the primary zone, greatly assists in breaking up
the fuel and mixing it with the incoming air.
Fig. 4.3 Flame stabilizing and general airflow pattern.
The temperature of the gases released by combustion is about 1,800 to 2,000 deg. C.,
which is far too hot for entry to the nozzle guide vanes of the turbine. The air not used for
combustion, which amounts to about 60 per cent of the total airflow, is therefore
introduced progressively into the flame tube. Approximately a third of this is used to
lower the gas temperature in the dilution zone before it entersthe turbine and the
remainder is used for cooling the walls of the flame tube. This is achieved by a film of
cooling air flowing along the inside surface of the flame tube wall, insulating it from the
hot combustion gases (fig. 4.4). A recent development allows cooling air to enter a
network of passages within the flame tube wall before exiting to form an insulating film
of air, this can reduce the required wall cooling airflow by up to 50 per cent. Combustion
should be completed before the dilution air enters the flame tube, otherwise the incoming
air will cool the flame and incomplete combustion will result.
Fig. 4.4 Flame tube cooling methods.
27
An electric spark from an igniter plug initiates combustion and the flame is then self-
sustained.
The design of a combustion chamber and the method of adding the fuel may vary
considerably, but the airflow distributions used to effect and maintain combustion is
always very similar to that described.
4.3 FUEL SUPPLY
Fuel is supplied to the airstream by one of two distinct methods. The most common is the
injection of a fine atomized spray into the recirculating airstream through spray nozzles.
The second method is based on the pre-vaporization of the fuel before it enters the
combustion zone.
Fig. 4.5 A vaporizer combustion chamber
In the vaporizing method (fig.4.5) the fuel is sprayed from feed tubes into vaporizing
tubes which are positioned inside the flame tube. These tubes turn the fuel through 180
degrees and, as they are heated by combustion, the fuel vaporizes before passing into the
flame tube. The primary airflow passes down the vaporizing tubes with the fuel and also
through holes in the flame tube entry section which provide 'fans' of air to sweep the
flame rearwards. Cooling and dilution air is metered into the flame tube in a manner
similar to the atomizer flame tube.
28
4.4 TYPES OF COMBUSTION CHAMBER
There are three main types of combustion chamber in use for gas turbine engines. These
are the multiple chambers, the tubo-annular chamber and the annular chamber.
Fig. 4.6 An early Whittle combustion chamber.
4.4.1 Annular combustion chamber
This type of combustion chamber consists of a single flame tube, completely annular in
form, which is contained in an inner and outer casing (fig. 4.7). The airflow through the
flame tube is similar to that already described, the chamber being open at the front to the
compressor and at the rear to the turbine nozzles.
The main advantage of the annular chamber is that, for the same power output, the length
of the chamber is only 75 per cent of that of a tubo-annular system of the same diameter,
resulting in considerable saving of weight and production cost. Another advantage is the
elimination of combustion propagation problems from chamber to chamber.
In comparison with a tubo-annular combustion system, the wall area of a comparable
annular chamber is much less; consequently the amount of cooling air required to prevent
the burning of the flame tube wall is less, by approximately 15 per cent, This reduction in
cooling air raises the combustion efficiency to virtually eliminate unburnt fuel, and
oxidizes the carbon monoxide to non-toxic carbon dioxide, thus reducing air pollution.
The main advantage of the annular chamber is that, for the same power output, the length
of the chamber is only 75 per cent of that of a tubo-annular system of the same diameter,
29
resulting in considerable saving of weight and production cost. Another advantage is the
elimination of combustion propagation problems from chamber to chamber.
Fig. 4.7 Annular combustion chamber.
In comparison with a tubo-annular combustion system, the wall area of a comparable
annular chamber is much less; consequently the amount of cooling air required to prevent
the burning of the flame tube wall is less, by approximately 15 per cent, This reduction in
cooling air raises the combustion efficiency to virtually eliminate unburnt fuel, and
oxidizes the carbon monoxide to non-toxic carbon dioxide, thus reducing air pollution.
The introduction of the air spray type fuel spray nozzle to this type of combustion
chamber also greatly improves the preparation of fuel for combustion by aerating the
over-rich pockets of fuel vapours close to the spray nozzle; this results in a large
reduction in initial carbon formation.
4.5 COMBUSTION CHAMBER PERFORMANCE
A combustion chamber must be capable of allowing fuel to burn efficiently over a wide
range of operating conditions without incurring a large pressure loss. In addition, if flame
extinction occurs, then it must be possible to relight. In performing these functions, the
flame tube and spray nozzle atomizer components must be mechanically reliable.
30
The gas turbine engine operates on a constant pressure cycle, therefore any loss of
pressure during the process of combustion must be kept to a minimum. In providing
adequate turbulence and mixing, a total pressure loss varying from about 3 to 8 per cent
of the air pressure at entry to the chamber is incurred.
4.5.1 Combustion intensity
The heat released by a combustion chamber or any other heat generating unit is
dependent on the volume of the combustion area. Thus, to obtain the required high power
output, a comparatively small and compact gas turbine combustion chamber must release
heat at exceptionally high rates.
For example, at take-off conditions a Rolls-Royce RB211-524 engine will consume
20,635 lb. of fuel per hour. The fuel has a calorific value of approximately 18,550 British
thermal units per lb., therefore the combustion chamber releases nearly 106,300 British
thermal units per second. Expressed in another way this is an expenditure of potential
heat at a rate equivalent to approximately 150,000 horsepower.
Fig. 4.8 Combustion efficiency and air/fuel Fig. 4.9 Combustion stability limits.
ratio.
4.5.2 Combustion efficiency
The combustion efficiency of most gas turbine engines at sea-level take-off conditions is
almost 100 per cent, reducing to 98 per cent at altitude cruise conditions, as shown in fig.
4.8.
4.5.3 Combustion stability
Combustion stability means smooth burning and the ability of the flame to remain alight
over a wide operating range.
For any particular type of combustion chamber there is both a rich and weak limit to the
air/fuel ratio, beyond which the flame is extinguished. An extinction is most likely to
occur in flight during a glide or dive with the engine idling, when there is a high airflow
and only a small fuel flow, i.e. a very weak mixture strength. Combustion stability means
smooth burning and the ability of the flame to remain alight over a wide operating range.
The range of air/fuel ratio between the rich and weak limits is reduced with an increase of
air velocity, and if the air mass flow is increased beyond a certain value, flame extinction
31
occurs. A typical stability loop is illustrated in fig. 4.9. The operating range defined by
the stability loop must obviously cover the air/fuel ratios and mass flow of the
combustion chamber.
The ignition process has weak and rich limits similar to those shown for stability in fig.
4.9. The ignition loop, however, lies within the stability loop since it is more difficult to
establish combustion under 'cold' conditions than to maintain normal burning.
4.5.6 Emissions
The unwanted pollutants which are found in the exhaust gases are created within the
combustion chamber. There are four main pollutants which are legislatively controlled;
unburnt hydrocarbons (unburnt fuel), smoke (carbon particles), carbon monoxide and
oxides of nitrogen. The principal conditions which affect the formation of pollutants are
pressure, temperature and time.
In the fuel rich regions of the primary zone, the hydrocarbons are converted into carbon
monoxide and smoke, fresh dilution air can be used to oxidize the carbon monoxide and
smoke into non-toxic carbon dioxide within the dilution zone. Unburnt hydrocarbons can
also be reduced in this zone by continuing the combustion process to ensure complete
combustion.
Oxides of nitrogen are formed under the same conditions as those required for the
suppression of the other pollutants, therefore it is desirable to cool the flame as quickly as
possible and to reduce the time available for combustion. This conflict of conditions
requires a compromise to be made, but continuing improvements in combustor design
and performance has led to a substantially 'cleaner' combustion process.
32
CHAPTER- 5
TURBINE
5.1 INTRODUCTION
The turbine has the task of providing the power to drive the compressor and accessories
and, in the case of engines which do not make use solely of a jet for propulsion, of
providing shaft power for a propeller or rotor. It does this by extracting energy from the
hot gases released from the combustion system and expanding them to a lower pressure
and temperature. High stresses are involved in this process, and for efficient operation,
the turbine blade tips may rotate at speeds over 1,500 feet per second, The continuous
flow of gas to which the turbine is exposed may have an entry temperature between 850
and 1,700 deg. C. and may reach a velocity of over 2,500 feet per second in parts of the
turbine.
Fig. 5.1 A triple-stage turbine with single shaft system.
To produce the driving torque, the turbine may consist of several stages each employing
one row of stationary nozzle guide vanes and one row of moving blades (fig. 5.1). The
number of stages depends upon the relationship between the power required from the gas
flow, the rotational speed at which it must be produced and the diameter of turbine
permitted.
33
Fig. 5.2 A twin turbine and shaft arrangement.
The number of shafts, and therefore turbines, varies with the type of engine; high
compression ratio engines usually have two shafts, driving high and low pressure
compressors (fig, 5.2). On high by-pass ratio fan engines that feature an intermediate
pressure system, another turbine may be interposed between the high and low pressure
turbines, thus forming a triple-spool system (fig, 5.3). On some engines, driving torque is
derived from a free-power turbine (fig. 5.4). This method allows the turbine to run at its
optimum speed because it is mechanically independent of other turbine and compressor
shafts.
The mean blade speed of a turbine has considerable effect on the maximum efficiency
possible for a given stage output. For a given output the gas velocities, deflections, and
hence losses, are reduced in proportion to the square of higher mean blade speeds. Stress
in the turbine disc increases as the square of the speed, therefore to maintain the same
stress level at higher speed the sectional thickness, hence the weight, must be increased
disproportionately. For this reason, the final design is a compromise between efficiency
and weight. Engines operating at higher turbine inlet temperatures are thermally more
efficient and have an improved power to weight ratio. By-pass engines have a better
propulsive efficiency and thus can have a smaller turbine for a given thrust.
34
Fig. 5.3 A triple turbine and shaft arrangement.
The design of the nozzle guide vane and turbine blade passages is based broadly on
aerodynamic considerations, and to obtain optimum efficiency, compatible with
compressor and combustion design, the nozzle guide vanes and turbine blades are of a
basic aerofoil shape. There are three types of turbine; impulse, reaction and a
combination of the two known as impulse-reaction. In the impulse type the total pressure
drop across each stage occurs in the fixed nozzle guide vanes which, because of their
convergent shape, increase the gas velocity whilst reducing the pressure. The gas is
directed onto the turbine blades which experience an impulse force caused by the impact
of the gas on the blades. In the reaction type the fixed nozzle guide vanes are designed to
alter the gas flow direction without changing the pressure. The converging blade passages
experience a reaction force resulting from the expansion and acceleration of the gas.
Normally gas turbine engines do not use pure impulse or pure reaction turbine blades but
the impulse-reaction combination (fig. 5.5). The proportion of each principle incorporated
in the design of a turbine is largely dependent on the type of engine in which the turbine
35
is to operate, but in general it is about 50 per cent impulse and 50 per cent reaction.
Impulse-type turbines are used for cartridge and air starters.
Fig. 5-4 A typical free power turbine.
5.2 ENERGY TRANSFER FROM GAS FLOW TO TURBINE
It will be seen that the turbine depends for its operation on the transfer of energy between
the combustion gases and the turbine. This transfer is never 100 per cent because of
thermodynamic and mechanical losses. When the gas is expanded by the combustion
process, it forces its way into the discharge nozzles of the turbine where, because of their
convergent shape, it is accelerated to about the speed of sound which, at the gas
temperature, is about 2,500 feet per second. At the same time the gas flow is given a 'spin'
or 'whirl' in the direction of rotation of the turbine blades by the nozzle guide vanes. On
impact with the blades and during the subsequent reaction through the blades, energy is
absorbed, causing the turbine to rotate at high speed and so provide the power for driving
the turbine shaft and compressor.
The torque or turning power applied to the turbine is governed by the rate of gas flow and
the energy change of the gas between the inlet and the outlet of the turbine blades, The
design of the turbine is such that the whirl will be removed from the gas stream so that
the flow at exit from the turbine will be substantially 'straightened out' to give an axial
36
flow into the exhaust system. Excessive residual whirl reduces the efficiency of the
exhaust system and also tends to produce jet pipe vibration which has a detrimental effect
on the exhaust cone supports and struts.
Fig. 5.5 Comparison between a pure Impulse turbine and an impulse/reaction turbine.
It will be seen that the nozzle guide vanes and blades of the turbine are 'twisted', the
blades having a stagger angle that is greater at the tip than at the root (fig. 5.6). The
reason for the twist is to make the gas flow from the combustion system do equal work at
all positions along the length of the blade and to ensure that the flow enters the exhaust
system with a uniform axial velocity. This results in certain changes in velocity, pressure
and temperature occurring through the turbine, as shown diagrammatically in fig. 5.7.
The 'degree of reaction' varies from root to tip, being least at the root and highest at the
tip, with the mean section having the chosen value of about 50 per cent.
The losses which prevent the turbine from being 100 per cent efficient are due to a
number of reasons. A typical uncooled three-stage turbine would suffer a 3.5 per cent loss
because of aerodynamic losses in the turbine blades. A further 4.5 per cent loss would be
incurred by aerodynamic losses in the nozzle guide vanes, gas leakage over the turbine
blade tips and exhaust system losses; these losses are of approximately equal proportions.
The total losses result in an overall efficiency of approximately 92 per cent. This results
in certain changes in velocity, pressure and temperature occurring through the turbine, as
shown diagrammatically in fig. 5.7
5.3 CONSTRUCTION
The basic components of the turbine are the combustion discharge nozzles, the nozzle
guide vanes, the turbine discs and the turbine blades. The rotating assembly is carried on
bearings mounted in the turbine casing and the turbine shaft may be common to the
compressor shaft or connected to it by a self-aligning coupling.
37
Fig. 5.6 A typical turbine blade showing twisted contour.
5.3.1 Nozzle guide vanes
The nozzle guide vanes are of an aerofoil shape with the passage between adjacent vanes
forming a convergent duct. The vanes are located (fig. 5.8) in the turbine casing in a
manner that allows for expansion.
Fig. 5.7 Gas flow pattern through nozzle and blade.
The nozzle guide vanes are usually of hollow form and may be cooled by passing
compressor delivery air through them to reduce the effects of high thermal stresses and
gas loads.
Turbine discs are usually manufactured from a machined forging with an integral shaft or
with a flange onto which the shaft may be bolted. The disc also has, around its perimeter,
provision for the attachment of the turbine blades.
To limit the effect of heat conduction from the turbine blades to the disc a flow of cooling
air is passed across both sides of each disc.
38
Fig. 5.8 Typical nozzle guide vanes showing their shape and location.
5.3.2 Turbine blades
The turbine blades are of an aerofoil shape, designed to provide passages between
adjacent blades that give a steady acceleration of the flow up to the 'throat', where the
area is smallest and the velocity reaches that required at exit to produce the required
degree of reaction.
The actual area of each blade cross-section is fixed by the permitted stress in the material
used and by the size of any holes which may be required for cooling purposes. High
efficiency demands thin trailing edges to the sections, but a compromise has to be made
so as to prevent the blades cracking due to the temperature changes during engine
operation.
The method of attaching the blades to the turbine disc is of considerable importance,
since the stress in the disc around the fixing or in the blade root has an important bearing
on the limiting rim speed. The blades on the early Whittle engine were attached by the de
Laval bulb root fixing, but this design was soon superseded by the 'fir-tree' fixing that is
now used in the majority of gas turbine engines. This type of fixing involves very
accurate machining to ensure that the loading is shared by all the serrations. The blade is
free in the serrations when the turbine is stationary and is stiffened in the root by
centrifugal loading when the turbine is rotating. Various methods of blade attachment are
shown in fig. 5.9; however, the B.M.W. hollow blade and the de Laval bulb root types
are not now generally used on gas turbine engines.
39
Fig. 5.9 Various methods of attaching blades to turbine discs.
A gap exists between the blade tips and casing, which varies in size due to the different
rates of expansion and contraction. To reduce the loss of efficiency through gas leakage
across the blade tips, a shroud is often fitted as shown in fig. 5.1. This is made up by a
small segment at the tip of each blade which forms a peripheral ring around the blade
tips. An abradable lining in the casing may also be used to reduce gas leakag. Active
Clearance Control (A.C.C.) is a more effective method of maintaining minimum tip
clearance throughout the flight cycle. Air from the compressor is used to cool the turbine
casing and when used with shroudless turbine blades, enables higher temperatures and
speeds to be used.
5.4 MATERIALS Among the obstacles in the way of using higher turbine entry temperatures have always
been the effects of these temperatures on the nozzle guide vanes and turbine blades, the
high speed of rotation which imparts tensile stress to the turbine disc and blades is also a
limiting factor.
5.4.1 Nozzle guide vanes
Due to their static condition, the nozzle guide vanes do not endure the same rotational
stresses as the turbine blades. Therefore, heat resistance is the property most required.
Nickel alloys are used, although cooling is required to prevent melting. Ceramic coatings
can enhance the heat resisting properties and, for the same set of conditions, reduce the
amount of cooling air required, thus improving engine efficiency.
40
5.4.2 Turbine discs
A turbine disc has to rotate at high speed in a relatively cool environment and is subjected
to large rotational stresses. The limiting factor which affects the useful disc life is its
resistance to fatigue cracking
Fig. 5.10 Free power contra-rotating turbine.
5.4.3 Turbine blades
A brief mention of some of the points to be considered in connection with turbine blade
design will give an idea of the importance of the correct choice of blade material. The
blades, while glowing red-hot, must be strong enough to carry the centrifugal loads due to
rotation at high speed. A small turbine blade weighing only two ounces may exert a load
of over two tons at top speed and it must withstand the high bending loads applied by the
gas to produce the many thousands of turbine horse-power necessary to drive the
compressor. Turbine blades must also be resistant to fatigue and thermal shock, so that
they will not fail under the influence of high frequency fluctuations in the gas conditions,
and they must also be resistant to corrosion and oxidization. In spite of all these demands,
the blades must be made in a material that can be accurately formed and machined by
current manufacturing methods.
From the foregoing, it follows that for a particular blade material and an acceptable safe
life there is an associated maximum permissible turbine entry temperature and a
corresponding maximum engine power. It is not surprising, therefore, that metallurgists
and designers are constantly searching for better turbine blade materials and improved
methods of blade cooling.
41
Over a period of operational time the turbine blades slowly grow in length. This
phenomenon is known as 'creep' and there is a finite useful life limit before failure occurs.
Fig. 5.11 Section through a dual alloy disc.
The early materials used were high temperature steel forgings, but these were rapidly
replaced by cast nickel base alloys which give better creep and fatigue properties.
Close examination of a conventional turbine blade reveals a myriad of crystals that lie in
all directions (equiaxed). Improved service life can be obtained by aligning the crystals to
form columns along the blade length, produced by a method known as 'Directional
Solidification'. A further advance of this technique is to make the blade out of a single
crystal.
A non-metal based turbine blade can be manufactured from reinforced ceramics. Their
initial production application is likely to be for small high speed turbines which have very
high turbine entry temperatures.
42
CHAPTER- 6
AFTERBURNING
6.1 INTRODUCTION
Afterburning (or reheat) is a method of augmenting the basic thrust of an engine to
improve the aircraft take-off, climb and (for military aircraft) combat performance. The
increased power could be obtained by the use of a larger engine, but as this would
increase the weight, frontal area and overall fuel consumption, afterburning provides the
best method of thrust augmentation for short periods. Afterburning consists of the
introduction and burning of fuel between the engine turbine and the jet pipe propelling
nozzle, utilizing the unburned oxygen in the exhaust gas to support combustion (fig. 6.1).
The resultant increase in the temperature of the exhaust gas gives an increased velocity of
the jet leaving the propelling nozzle and therefore increases the engine thrust.
Fig. 6.1 Principle of afterburning
As the temperature of the afterburner flame can be in excess of 1,700 deg. C., the burners
are usually arranged so that the flame is concentrated around the axis of the jet pipe. This
allows a proportion of the turbine discharge gas to flow along the wall of the jet pipe and
thus maintain the wall temperature at a safe value.
The area of the afterburning jet pipe is larger than a normal jet pipe would be for the
same engine, to obtain a reduced velocity gas stream. To provide for operation under all
conditions, an afterburning jet pipe is fitted with either a two-position or a variable-area
propelling nozzle (fig. 6.2) the nozzle is closed during non-afterburning operation, but
when afterburning is selected the gas temperature increases and the nozzle opens to give
an exit area suitable for the resultant increase in the volume of the gas stream. This
prevents any increase in pressure occurring in the jet pipe which would affect the
functioning of the engine and enables afterburning to be used over a wide range of engine
speeds.
The thrust of an afterburning engine, without afterburning in operation, is slightly less
than that of a similar engine not fitted with afterburning equipment; this is due to the
43
added restrictions in the jet pipe. The overall weight of the power plant is also increased
because of the heavier jet pipe and after-burning equipment.
Fig. 6.2 Examples of afterburning jet pipes and propelling nozzles.
Afterburning is achieved on low by-pass engines by mixing the by-pass and turbine
streams before the afterburner fuel injection and stabilizer system is reached so that the
combustion takes place in the mixed exhaust stream. An alternative method is to inject
the fuel and stabilize the flame in the individual by-pass and turbine streams, burning the
available gases up to a common exit temperature at the final nozzle. In this method, the
fuel injection is scheduled separately to the individual streams and it is normal to provide
some form of interconnection between the flame stabilizers in the hot and cold streams to
assist the combustion processes in the cold by-pass air. In this method, the fuel injection
is scheduled separately to the individual streams and it is normal to provide some form of
44
interconnection between the flame stabilizers in the hot and cold streams to assist the
combustion processes in the cold by-pass air.
6.2 OPERATION OF AFTERBURNING
The gas stream from the engine turbine enters the jet pipe at a velocity of 750 to 1,200
feet per second, but as this velocity is far too high for a stable flame to be maintained, the
flow is diffused before it enters the afterburner combustion zone, i.e. the flow velocity is
reduced and the pressure is increased. However, as the speed of burning kerosene at
normal mixture ratios is only a few feet per second, any fuel lit even in the diffused air
stream would be blown away. A form of flame stabilizer (vapour gutter) is, therefore,
located downstream of the fuel burners to provide a region in which turbulent eddies are
formed to assist combustion and where the local gas velocity is further reduced to a figure
at which flame stabilization occurs whilst combustion is in operation.
Fig. 6.3 Methods of afterburning ignition.
An atomized fuel spray is fed into the jet pipe through a number of burners, which are so
arranged as to distribute the fuel evenly over the flame area. Combustion is then initiated
by a catalytic igniter, which creates a flame as a result of the chemical reaction of the
fuel/air mixture being sprayed on to a platinum-based element, by an igniter plug
adjacent to the burner, or by a hot streak of flame that originates in the engine combustion
chamber (fig 6.3): this latter method is known as 'hot-shot' ignition. Once combustion is
initiated, the gas temperature increases and the expanding gases accelerate through the
enlarged area propelling nozzle to provide the additional thrust.
45
In view of the high temperature of the gases entering the jet pipe from the turbine, it
might be assumed that the mixture would ignite spontaneously. This is not so, for
although cool flames form at temperatures up to 700 deg. C., combustion will not take
place below 800 deg. C. If however, the conditions were such that spontaneous ignition
could be effected at sea level; it is unlikely that it could be effected at altitude where the
atmospheric pressure is low. The spark or flame that initiates combustion must be of such
intensity that a light-up can be obtained at considerable altitudes.
For smooth functioning of the system, a stable flame that will burn steadily over a wide
range of mixture strengths and gas flows is required. The mixture must also be easy to
ignite under all conditions of flight and combustion must be maintained with the
minimum loss of pressure.
6.3 CONSTRUCTION
6.3.1 Burners
The burner system consists of several circular concentric fuel manifolds supported by
struts inside the jet pipe. Fuel is supplied to the manifolds by feed pipes in the support
struts and sprayed into the flame area, between the flame stabilizers, from holes in the
downstream edge of the manifolds. The flame stabilizers are blunt nosed V-section
annular rings located downstream of the fuel burners. An alternative system includes an
additional segmented fuel manifold mounted within the flame stabilizers. The typical
burner and flame stabilizer shown in fig. 6.4 is based on the latter system.
6.3.2 Jet pipe
The afterburning jet pipe is made from a heat-resistant nickel alloy and requires more
insulation than the normal jet pipe to prevent the heat of combustion being transferred to
the aircraft structure. The jet pipe may be of a double skin construction with the outer
skin carrying the flight loads and the inner skin the thermal stresses; a flow of cooling air
is often induced between the inner and outer skins. Provision is also made to
accommodate expansion and contraction, and to prevent gas leaks at the jet pipe joints.
A circular heatshield of similar material to the jet pipe is often fitted to the inner wall of
the jet pipe to improve cooling at the rear of the burner section. The heatshield comprises
a number of bands, linked by cooling corrugations, to form a single skin. The rear of the
heatshield is a series of overlapping 'tiles' riveted to the surrounding skin (fig. 6.4). The
shield also prevents combustion instability from creating excessive noise and vibration,
which in turn would cause rapid physical deterioration of the afterburner equipment.
6.3.3 Propelling nozzle
The propelling nozzle is of similar material and construction as the jet pipe, to which it is
secured as a separate assembly. A two-position propelling nozzle has two movable
eyelids that are operated by actuators, or pneumatic rams, to give an open or closed
position. A variable-area propelling nozzle has a ring of interlocking flaps that are hinged
to the outer casing and may be enclosed by an outer shroud. The flaps are actuated by
powered rams to the closed position, and by gas loads to the intermediate or the open
46
positions; control of the flap position is by a control unit and a pump provides the power
to the rams.
Fig. 6.4 Typical afterburning jet pipe equipment.
6.4 CONTROL SYSTEM
It is apparent that two functions, fuel flow and propelling nozzle area, must be
coordinated for satisfactory operation of the afterburner system, these functions are
related by making the nozzle area dependent upon the fuel flow at the burners or vice-
versa. The pilot controls the afterburner fuel flow or the nozzle area in conjunction with a
compressor delivery/jet pipe pressure sensing device (a pressure ratio control unit). When
the afterburner fuel flow is increased, the nozzle area increases; when the afterburner fuel
flow decreases, the nozzle area is reduced. The pressure ratio control unit ensures the
pressure ratio across the turbine remains unchanged and that the engine is unaffected by
the operation of afterburning, regardless of the nozzle area and fuel flow. Since large fuel
flows are required for afterburning, an additional fuel pump is used. This pump is usually
of the centrifugal flow or gear type and is energized automatically when afterburning is
selected. The system is fully automatic and incorporates 'fail safe' features in the event of
an afterburner malfunction. The interconnection between the control system and
afterburner jet pipe is shown diagrammatically in fig. 6.5.
47
Fig. 6.5 Simplified control system.
When afterburning is selected, a signal is relayed to the afterburner fuel control unit. The
unit determines the total fuel delivery of the pump and controls the distribution of fuel
flow to the burner assembly. Fuel from the burners is ignited, resulting in an increase in
jet pipe pressure (P6). This alters the pressure ratio across the turbine (P3/P6), and the
exit area of the jet pipe nozzle is automatically increased until the correct PS/PS ratio has
been restored. With a further increase in the degree of afterburning, the nozzle area is
progressively increased to maintain a satisfactory P3/P6 ratio. Fig. 6.6 illustrates a typical
afterburner fuel control system.
To operate the propelling nozzle against the large 'drag' loads imposed by the gas stream,
a pump and either hydraulically or pneumatically operated rams are incorporated in the
control system. The system shown in fig. 6.7 uses oil as the hydraulic medium, but some
systems use fuel. Nozzle movement is achieved by the hydraulic operating rams which
are pressurized by an oil pump, pump output being controlled by a linkage from the
pressure ratio control unit.
When an increase in afterburning is selected, the afterburner fuel control unit schedules
an increase in fuel pump output. The jet pipe pressure (P6) increases, altering the
pressure ratio across the turbine (P3/P6).
The pressure ratio control unit alters oil pump output, causing an out-of-balance
condition between the hydraulic ram load and the gas load on the nozzle flaps.
48
Fig. 6.6 A simplified typical afterburner fuel control system.
The gas load opens the nozzle to increase its exit area and, as the nozzle opens, the
increase in nozzle area restores the P3/P6 ratio and the pressure ratio control unit alters
oil pump output until balance is restored between the hydraulic rams and the gas loading
on the nozzle flaps.
When afterburning is selected, a signal is relayed to the afterburner fuel control unit. The
unit determines the total fuel delivery of the pump and controls the distribution of fuel
flow to the burner assembly. Fuel from the burners is ignited, resulting in an increase in
jet pipe pressure (P6). This alters the pressure ratio across the turbine (P3/P6), and the
49
exit area of the jet pipe nozzle is automatically increased until the correct PS/PS ratio has
been restored. With a further increase in the degree of afterburning, the nozzle area is
progressively increased to maintain a satisfactory P3/P6 ratio. Fig. 6.6 illustrates a typical
afterburner fuel control system.
Fig. 6.7 A simplified typical afterburner nozzle control system.
6.5 THRUST INCREASE
The increase in thrust due to afterburning depends solely upon the ratio of the absolute jet
pipe temperatures before and after the extra fuel is burnt. For example, neglecting small
losses due to the afterburner equipment and gas flow momentum changes, the thrust
increase may be calculated as follows.
Assuming a gas temperature before afterburning of 640 deg. C and with afterburning of
1,269 deg. C. (1,542 deg. K.) then the temperature ratio = 1,542 = 1.69. 913.
The velocity of the jet stream increases as the square root of the temperature ratio.
Therefore, the jet velocity = ^/T.69 = 1.3. Thus, the jet stream velocity is increased by 30
per cent, and the increase in static thrust, in this instance, is also 30 per cent (fig. 6.8).
The increase in thrust due to afterburning depends solely upon the ratio of the absolute jet
pipe temperatures before and after the extra fuel is burnt. For example, neglecting small
50
losses due to the afterburner equipment and gas flow momentum changes, the thrust
increase may be calculated as follows.
Static thrust increases of up to 70 per cent are obtainable from low by-pass engines fitted
with after-burning equipment and at high forward speeds several times this amount of
thrust boost can be obtained. High thrust boosts can be achieved on low by-pass engines
because of the large amount of oxygen in the exhaust gas stream and the low initial
temperature of the exhaust gases.
It is not possible to go on increasing the amount of fuel that is burnt in the jet pipe so that
all the available oxygen is used, because the jet pipe would not withstand the high
temperatures that would be incurred and complete combustion cannot be assured.
Fig. 6.8 Thrust increase and temperature Fig. 6.9 Specific fuel consumption
ratio. comparison.
6.6 FUEL CONSUMPTION
Afterburning always incurs an increase in specific fuel consumption and is, therefore,
generally limited to periods of short duration. Additional fuel must be added to the gas
stream to obtain the required temperature ratio. Since the temperature rise does not occur
at the peak of compression, the fuel is not burnt as efficiently as in the engine combustion
chamber and a higher specific fuel consumption must result. For example, assuming a
specific fuel consumption without after-burning of 1,15 lb./hr./lb. thrust at sea level and a
speed of Mach 0,9 as shown in fig. 6.9. then with 70 per cent afterburning under the same
conditions of flight, the consumption will be increased to approximately 2.53 lb./hr./lb.
thrust. With an increase in height to 35,000 feet this latter figure of 2.53 lb./hr./lb. thrust
will fall slightly to about 2.34 lb./hr./lb. thrust due to the reduced intake temperature.
51
Fig. 6.10 Afterburning and its effect on the rate of climb.
When this additional fuel consumption is combined with the improved rate of take-off
and climb (fig. 6.10), it is found that the amount of fuel required to reduce the time taken
to reach operation height is not excessive.
52
CHAPTER- 7
EXHAUST SYSTEM
7.1 INTRODUCTION
Aero gas turbine engines have an exhaust system which passes the turbine discharge
gases to atmosphere at a velocity, and in the required direction, to provide the resultant
thrust. The velocity and pressure of the exhaust gases create the thrust in the turbo-jet
engine but in the turbo-propeller engine only a small amount of thrust is contributed by
the exhaust gases, because most of the energy has been absorbed by the turbine for
driving the propeller. The design of the exhaust system therefore, exerts a considerable
influence on the performance of the engine. The areas of the jet pipe and propelling or
outlet nozzle affect the turbine entry temperature, the mass airflow and the velocity and
pressure of the exhaust jet.
Fig. 7.1 A basic exhaust system.
The temperature of the gas entering the exhaust system is between 550 and 850 deg. C.
according to the type of engine and with the use of afterburning can be 1,500 deg. C. or
higher. Therefore, it is necessary to use materials and a form of construction that will
resist distortion and cracking, and prevent heat conduction to the aircraft structure.
A basic exhaust system is shown in fig. 7.1. The use of a thrust reverser, noise suppressor
and a two position propelling nozzle entails a more complicated system as shown in fig.
7.2. The low by-pass engine may also include a mixer unit (fig. 7.4) to encourage a
thorough mixing of the hot and cold gas streams.
53
Fig. 7.2 Exhaust system with thrust reverser, noise suppressor and two position
propelling nozzle.
7.2 EXHAUST GAS FLOW
Gas from the engine turbine enters the exhaust system at velocities from 750 to 1,200 feet
per second, but, because velocities of this order produce high friction losses, the speed of
flow is decreased by diffusion. This is accomplished by having an increasing passage
area between the exhaust cone and the outer wall as shown in fig. 7.1. The cone also
prevents the exhaust gases from flowing across the rear face of the turbine disc. It is usual
to hold the velocity at the exhaust unit outlet to a Mach number of about 0.5, i.e.
approximately 950 feet per second. Additional losses occur due to the residual whirl
velocity in the gas stream from the turbine. To reduce these losses, the turbine rear struts
in the exhaust unit are designed to straighten out the flow before the gases pass into the
jet pipe.
The exhaust gases pass to atmosphere through the propelling nozzle, which is a
convergent duct, thus increasing the gas velocity. In a turbo-jet engine, the exit velocity
of the exhaust gases is subsonic at low thrust conditions only. During most operating
conditions, the exit velocity reaches the speed of sound in relation to the exhaust gas
temperature and the propelling nozzle is then said to be 'choked'; that is, no further
increase in velocity can be obtained unless the temperature is increased. As the upstream
total pressure is increased above the value at which the propelling nozzle becomes
'choked', the static pressure of the gases at exit increases above atmospheric pressure.
This pressure difference across the propelling nozzle gives what is known as 'pressure
thrust' and is effective over the nozzle exit area. This is additional thrust to that obtained
due to the momentum change of the gas stream.
54
Fig. 7.3 Gas flow through a convergentdivergent nozzle.
With the convergent type of nozzle, wastage of energy occurs, since the gases leaving the
exit do not expand rapidly enough to immediately achieve outside air pressure.
Depending on the aircraft flight plan, some high pressure ratio engines can with
advantage use a convergent-divergent nozzle to recover some of the wasted energy this
nozzle utilizes the pressure energy to obtain a further increase in gas velocity and,
consequently, an increase in thrust.
Fig. 7.4 A low by-pass air mixer unit.
55
From the illustration (fig. 7.3), it will be seen that the convergent section exit now
becomes the throat, with the exit proper now being at the end of the flared divergent
section. When the gas enters the convergent section of the nozzle, the gas velocity
increases with a corresponding fall in static pressure. The gas velocity at the throat
corresponds to the local sonic velocity. As the gas leaves the restriction of the throat and
flows into the divergent section, it progressively increases in velocity towards the exit.
The reaction to this further increase in momentum is a pressure force acting on the inner
wall of the nozzle. A component of this force acting parallel to the longitudinal axis of
the nozzle produces the further increase in thrust.
Fig. 7.5 High by-pass ratio engine exhaust systems.
56
The propelling nozzle size is extremely important and must be designed to obtain the
correct balance of pressure, temperature and thrust. With a small nozzle these values
increase, but there is a possibility of the engine surging, whereas with a large nozzle the
values obtained are too low,
A fixed area propelling nozzle is only efficient over a narrow range of engine operating
conditions. To increase this range, a variable area nozzle may be used. This type of
nozzle is usually automatically controlled and is designed to maintain the correct balance
of pressure and temperature at all operating conditions. In practice, this system is seldom
used as the performance gain is offset by the increase in weight. However, with
afterburning a variable area nozzle is necessary.
The by-pass engine has two gas streams to eject to atmosphere, the cool by-pass airflow
and the hot turbine discharge gases.
In a low by-pass ratio engine, the two flows are combined by a mixer unit (fig. 7.4) which
allows the by-pass air to flow into the turbine exhaust gas flow in a manner that ensures
thorough mixing of the two streams.
In high by-pass ratio engines, the two streams are usually exhausted separately. The hot
and cold nozzles are co-axial and the area of each nozzle is designed to obtain maximum
efficiency. However, an improvement can be made by combining the two gas flows
within a common, or integrated, nozzle assembly. This partially mixes the gas flows prior
to ejection to atmosphere. An example of both types of high by-pass exhaust system is
shown in fig, 7.5.
7.3 CONSTRUCTION AND MATERIALS
The exhaust system must be capable of with-standing the high gas temperatures and is
therefore manufactured from nickel or titanium. It is also necessary to prevent any heat
being transferred to the surrounding aircraft structure. This is achieved by passing
ventilating air around the jet pipe, or by lagging the section of the exhaust system with an
insulating blanket (fig. 7.6). Each blanket has an inner layer of fibrous insulating material
contained by an outer skin of thin stainless steel, which is dimpled to increase its strength.
In addition, acoustically absorbent materials are sometimes applied to the exhaust system
to reduce engine noise.
Fig. 7.6 An insulating blanket.
When the gas temperature is very high (for example, when afterburning is employed), the
complete jet pipe is usually of double-wall construction with an annular space between
the two walls. The hot gases leaving the propelling nozzle induce, by ejector action, a
flow of air through the annular space of the engine nacelle. This flow of air cools the
57
inner wall of the jet pipe and acts as an insulating blanket by reducing the transfer of heat
from the inner to the outer wall.
The cone and streamline fairings in the exhaust unit are subjected to the pressure of the
exhaust gases; therefore, to prevent any distortion, vent holes are provided to obtain a
pressure balance.
The mixer unit used in low by-pass ratio engines consists of a number of chutes through
which the bypass air flows into the exhaust gases. A bonded honeycomb structure is used
for the integrated nozzle assembly of high by-pass ratio engines to give lightweight
strength to this large component.
Due to the wide variations of temperature to which the exhaust system is subjected, it
must be mounted and have its sections joined together in such a manner as to allow for
expansion and contraction without distortion or damage.
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CONCLUSION
The Jet engine's invention changed the future. With jet engines, planes can carry more
cargo, fly faster, and go farther than any propeller plane.Today, the fastest passenger jet
flies from London to New York in 7-8 hours. Planes have made it easier to see new
places, and experience new cultures. Billions of people have flown in airplanes and the
number keeps getting bigger, so it's safe to say that it's changed these peoples lives, and
has changed the world.
Wars are also fought with jets. In the Iraq war we've used jets to fight terrorism, and also
in Afghanistan. Plus, old battleships have been converted to aircraft carriers because
fighter jets are simply better. Jet engine designs are frequently modified for non-aircraft
applications, as industrial gas turbines or marine power plants. These are used in
electrical power generation, for powering water, natural gas, or oil pumps, and providing
propulsion for ships and locomotives. Industrial gas turbines can create up to 50,000 shaft
horsepower. Many of these engines are derived from older military turbojets such as the
Pratt & Whitney J57 and J75 models. There is also a derivative of the P&W JT8D low-
bypass turbofan that creates up to 35,000 HP. Jet engines are also sometimes developed
into, or share certain components such as engine cores, with turboshaft and turboprop
engines, which are forms of gas turbine engines that are typically used to power
helicopters and some propeller-driven aircraft.
Jets have changed the way people lived. They've made traveling faster, and more
efficient, and have made a big change in militaries all over the world. The jet engine
changed history.
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REFERENCES
[1]. https://en.m.wikipedia.org/wiki/Jet_engine
[2]. http://www.explainthatstuff.com/jetengine.html
[3]. http://engineering.mit.edu/ask/how-does-jet-engine-work
[4]. https://blog.klm.com/8-things-you-probably-dont-know-about-jet-engines/
[5]. https://www.grc.nasa.gov/www/k-12/UEET/StudentSite/engines.html
[6]. https://animagraffs.com/inside-a-jet-engine/
[7]. http://www.pbsvb.com/customer-industries/aerospace/aircraft-engines/tj-100-turbojet-engine
[8]. http://www.historylearningsite.co.uk/inventions-and-discoveries-of-the-twentieth-
century/the-jet-engine/
[9]. https://www.geaviation.com/company/aviation-history
[10]. http://science.howstuffworks.com/transport/flight/modern/turbine.html
[11]. http://www.bloodhoundssc.com/project/car/engines/jet-engine
[12]. http://books.google.co.in/books/about/The_Jet_Engine.html?id=jv5ZAAAAYAAJ
[13].http://airspot.ru/book/file/485/166837_EB161_rolls_royce_the_jet_engine_fifth_edition_
gazoturbinnyy_dviga.pdf
[14]. http://www.barnesandnoble.com/mobile/w/jet-engines-klaus-hunecke/1012585499
60
APPENDIX
in = inch
ft = foot
yd = yard
oz = ounce
lb = pound
cwt = hundredweight
Btu = British thermal unit
hp = horsepower
Hg = mercury
s = second
min = minute
h = hour
f = force
W = watt
kW = kilowatt (Wx1000)
mm = millimetre (mx0.001)
m = metre
km = kilometre (mx1000)
g = gram
kg = kilogram
N = newton
Pa = pascal
kPa = kilopascal
J = Joule
kJ = kilojoule (Jx1000)
MJ = megajoule (Jx1 000 000)